The present invention relates to a fuel jetting nozzle assembly for use in a gas turbine combustor, particularly, in which a burn damage to the central portion of the extreme end portion of the fuel jetting nozzle is prevented as much as possible.
FIGS. 9 to 11, explained hereinafter, represents an example of a typical well-known gas turbine combustor of a conventional structure.
Referring to FIGS. 9 to 11, a plurality of gas turbine combustors are arranged on the outer peripheral portion of a discharge casing 2 of an air conditioner 1. A combustor liner 5 by which an internal combustion chamber 4 is enclosed is housed within the combustor casing 3, and a nozzle head 6, an igniter 7 and a flame detector, not shown, are provided in the internal combustion chamber 4. The nozzle head 6 is mounted on a head plate 8, and this head plate 8 and a flow sleeve 9 are mounted on the combustor casing 3. The head plate 8 is disposed so as to close one end of the casing 3.
A fuel jetting nozzle 10 is mounted on the nozzle head 6 and prevented from rotating by a locking plate 11. The combustor liner 5 is mounted on the extreme, i.e. front, end portion of the fuel jetting nozzle 10, and a liner supporter 12 provided on the flow sleeve 9 supports the combustor liner 5.
A transition piece 13 is connected to the extreme end portion (the downstream area) of the combustor liner 5. The combustor liner 5 is connected to a first-stage turbine stationary blade 14a of a gas turbine 14 by way of the transition piece 13.
An air intake passage 15 is formed in the outer peripheral portion of the fuel jetting nozzle 10. A swirl vane 16 is disposed between the air intake passage 15 and the internal combustion chamber 4. Fuel jetting holes 17, through which the inside of the fuel jetting nozzle 10 is communicated with the swirl vane 16, are provided on the peripheral wall portion of the fuel jetting nozzle 10.
The front side of a central end portion 18 of the fuel jetting nozzle 10 faces the inside of the internal combustion chamber 4 and forms a portion thereof. A fuel intake 19 is formed in the nozzle head 6, from which a gaseous fuel 20 is introduced into the fuel jetting nozzle 10.
An air flow around the gas turbine combustor will be explained hereunder.
An air 21 discharged from the air conditioner 1 flows around the transition piece 13 and is guided in a direction opposite to the flow of combustion gas 22 between the combustor liner 5 and the flow sleeve 9. The discharged air 21 is introduced into the internal combustion chamber 4 through air passages which are broadly divided into three portions. That is, the discharged air 21 is divided into primary air 23 introduced from the swirl vane 16 around the fuel jetting nozzle 10, secondary air 25 introduced from an air guide 24 provided on the trunk portion of the combustor liner 5, and tertiary air 26 for dilution purposes introduced from the holes provided downstream of the air guide 24 used for the secondary air.
A stable annular vortex area, i.e. flame area, of the primary air 23 and the gaseous fuel 20 is formed in the inside of the annular swirl flow caused by the primary air 23. The stable annular vortex area stabilizes and maintains the combustion flame, and the combustion gas 22 flows to the exit area of the combustor liner 5. The primary air 23 is mixed with the tertiary air 26, and cools the combustor liner 5 and decreases the gas temperature so that the liner exit temperature becomes a temperature required for the turbine.
In this viewpoint, the primary air 23, the secondary air 25 and the tertiary air 26 are allocated in various ways so as to control combustion performance. In some instances, the secondary air 25 and tertiary air 26 may not be provided. Furthermore, the primary air 23 and the secondary air 25 may be mixed with the gaseous fuel 20 beforehand and introduced into the internal combustion chamber 4.
The discharged air 21 passes through a slot, not shown, used to cool the combustor liner 5 and is supplied to the internal combustion chamber 4.
The details of the fuel jetting nozzle 10 are shown in FIG. 10.
Some of the primary air 23 of the discharged air 21 discharged from the air compressor 1 enters from the air intake passage 15 into the internal combustion chamber 4. At this time, the air is mixed with the gaseous fuel 20 jetted from the fuel jetting holes 17, passes the swirl vane 16 disposed around the fuel jetting nozzle 10, is jetted into the internal combustion chamber 4 while it is being swirled and is then ignited. Ignition is performed by the igniter 7 shown in FIG. 9. The combustion gas 22 passes the transition piece 13 and is introduced to the first-stage turbine stationary blade 14a of the gas turbine 14, causing a turbine rotor, not shown, to rotate by using the energy thereof.
The flow of gas near the outlet of the fuel jetting nozzle 10 inside the internal combustion chamber 4 is shown in FIG. 11.
The primary air 23 passes the swirl vanes 16 of the fuel jetting nozzle 10 and flows into the internal combustion chamber 4 while it is being swirled. The secondary air 25 which flows into the internal combustion chamber 4 through the air guide 24 provided in the trunk portion of the combustor liner 5 flows into a swirling flow 27 formed by air passing through the fuel jetting nozzle 10, forming a reverse flow, i.e. vortex flow, flame area 28 in the central portion and a reverse flow, i.e. vortex flow, flame area 29 in the outer periphery. The local temperature of the combustion gas inside the reverse flow flame area 28 in the central portion becomes a high temperature above approximately 2,000.degree. C. and a stable flame can thus be maintained.
However, in the above-described conventional fuel jetting nozzle for use in a gas turbine combustor, problems arise. For example, the central end portion 18 of the fuel jetting nozzle 10 is burned by radiation and forced convection by high-temperature gas of the reverse flow flame area 28 in the central portion, and the service life of the fuel jetting nozzle 10 becomes short.